1. Field of the Invention
This invention relates to bypass gas turbine engines, particularly means for varying gas flow through the discharge nozzle thereof.
2. The Prior Art
In a typical bypass gas turbine engine, a core duct having a compressor, combustion section, high-pressure turbine, low-pressure turbine, a plurality of outlet guide vanes and a discharge nozzle is surrounded by an annular bypass duct which communicates with the core duct by movable valves, e.g., aft or downstream of the outlet guide vane and before or upstream of the discharge nozzle to mix (fan) bypass air with core gas and by controlling the opening and closing of such bypass vanes, to vary the bypass ratio of such engine.
Thus when maximum power is desired, e.g., for an ascending aircraft, the duct vanes are virtually closed, to reduce the flow of by-pass air into the core and therethrough known as minimum bypass ratio. During cruise or part-power conditions, such duct vanes would be increasingly opened to increase the flow of by-pass air into the core, reducing engine power and thus have the engine operate at high bypass ratio.
However such bypass vanes have proved cumbersome in operation and have added weight and complexity to the engine.
A further disadvantage of prior art variable bypass ratio gas turbine engines is that they have failed to meet desired performance goals in all modes of operation because they have performed with insufficient flow variability to maintain satisfactory engine performance in both supersonic and subsonic flight.
Attempts have been made to improve the flow variability in the above bypass gas turbine engines, see for example U.S. Pat. No. 4,050,242 to Dusa (1977) and U.S. Pat. No. 4,069,661 to Rundell et al (1978). These references teach the axially shifting of humpbacked surfaces in the bypass annulus of the gas turbine engine to constrict the annular bypass area in such engine to vary the bypass ratio per FIG. 3 of the above 1978 reference and FIGS. 2 and 3 of the above 1977 reference. Alternatively, a series of louvers connecting bypass duct to core duct positioned downstream of the low pressure turbine of such engine are opened and closed to vary the ratio of bypass air injected into the core gas duct per FIGS. 1 and 2 of the above 1978 reference.
The above references, while directed to apparatus which can enlarge or constrict flow area between bypass duct and core duct have added components of complexity to such engines with attendant breakdown and maintenance problems and would add considerable weight to such engine.
The above references also do not address correcting the swirl of exit gases from the low-pressure turbine to flow axially toward the discharge nozzle.
There is, accordingly, a need and market for an engine component which can overcome the above prior art shortcomings in a) swirl correction and/or b) in variable area bypass injection.
There has now been discovered an apparatus for counteracting and/or redirecting turbine exit gas swirl. There has further been discovered an apparatus for varying the area of bypass gas injection from bypass duct to core duct in a gas turbine engine that is of reduced complexity and lighter in weight than available in the prior art.